Oxidizer-rich liquid monopropellants for a dual mode chemical rocket engine

ABSTRACT

The subject invention relates to oxidizer-rich liquid monopropellants based on ADN or HAN for a dual mode bipropellant chemical rocket engine. Such engines may be part of propulsion systems to be used in aerospace applications for 1) orbit raising, orbit manoeuvres and maintenance, attitude control and deorbiting of spacecraft, and/or 2) propellant settling, attitude and roll control of missiles, launchers and space planes.

FIELD OF THE INVENTION

The subject invention relates to oxidizer-rich liquid monopropellantsbased on ADN or HAN for a dual mode bipropellant chemical rocket engine.Such engines may be part of propulsion systems to be used in aerospaceapplications for 1) orbit raising, orbit manoeuvres and maintenance,attitude control and deorbiting of spacecraft, and/or 2) propellantsettling, attitude and roll control of missiles, launchers and spaceplanes.

BACKGROUND OF THE INVENTION

Dual mode rocket propulsion systems and dual mode rocket engines (alsoreferred to as thrusters) are known in the art. Currently, manyspacecraft use dual-mode propulsion systems, with bipropellant enginesfor larger thrust operations, and monopropellant engines for smallerthrust or when minimum impulse bit is important. In the art the choiceof propellants which are suitable in both bipropellant andmonopropellant engines are limited to a few very hazardous propellants.Such bipropellants comprise hydrazine or a derivative thereof, such asmonomethyl hydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH).An example of a dual mode thruster is a thruster referred to as aSecondary Combustion Augmented Thruster (SCAT). A bipropellant dual moderocket propulsion system comprising a bipropellant thruster having dualmode capability (i.e. ability to operate either in monopropellant modeor in bipropellant mode) has been described in e.g. U.S. Pat. No.6,135,393, wherein hydrazine is used as the fuel, and, preferably,nitrogen tetroxide (NTO) as the oxidizer.

The mission requirements for a particular propulsion system requiringhigh performance are defined by a set of figures of merit. One of themost important figures of merit is specific impulse (I_(sp)) as itindicates the maximum velocity changes that the spacecraft can achieve,which is the very objective of such propulsion system. Specific impulseis defined as the thrust developed by an engine per unit of propellantmass flow rate. If the thrust is measured in Newton (N) and the flowrate is measured in kilograms (kg) per second (s), then the unit ofmeasurement of specific impulse is Ns/kg. For medium to large spacecraftwith requirements of significant velocity changes this is the mostimportant parameter. For small spacecraft where dimensions may belimiting, the density impulse, i.e. Ns per propellant volume, may be thedominant figure of merit. Another figure of merit is the thrust of arocket engine as it determines how long a maneuver will take and whatacceleration it will provide. Yet another parameter is the smallest orminimum impulse bit (Ns) that the engine can generate as it determineshow precise a maneuver can be performed.

Both hydrazine (fuel) and nitrogen tetroxide (oxidizer), and theirderivatives are extremely hazardous for humans as they are highly toxic,carcinogenic, corrosive, etc., and they are associated with significantconcerns regarding the severe impact on the environment that they cancause in the case of spillage and emissions Therefore, the handlingthereof and the safety requirements are extremely demanding, timeconsuming and costly.

The ECHA (European Chemicals Agency) has within REACH (Registration,Evaluation, Authorisation and restriction of Chemicals), which is theEuropean Community Regulation on chemicals and their safe use,identified hydrazine as a substance of very high concern which may leadto that hydrazine may be banned for use in new development. Clean Space,which is an initiative by the European Space Agency (ESA), also callsfor substituting conventional hazardous propellants.

There is also a new law, Space Operations Act, in France, with respectto space debris, which requires that the spacecraft shall be deorbitedwhen no longer in use.

A significant achievement in the art is the feasibility to substitutehydrazine as a monopropellant for many space applications. This has beensuccessfully demonstrated using the HPGP® technology comprising theLMP-103S monopropellant blend (described in WO 2012/166046) andcorresponding thrusters (disclosed in e.g. WO 02/095207) ranging fromtypically 0.5 N to 200 N. A 1 N HPGP® propulsion system has beenoperational for several years in an earth orbit in space on the mainPRISMA satellite.

Accordingly, it is therefore desirable to provide a propellant enablinga dual mode propulsion system avoiding the use of hydrazine, nitrogentetroxide, and derivatives thereof However, so far, no viable rocketpropulsion systems, rocket engines, and corresponding alternativepropellants with performance comparable to the prior art hazardoushydrazine propellants have been realized.

SUMMARY OF THE INVENTION

The present inventors have developed a low-hazard oxidizer-rich liquidmonopropellant comprising 70-90% of an oxidizer selected from ADN andHAN; 0-10% ammonia, and; balance water, which can be used inbipropellant mode operation in a chemical rocket engine in combinationwith a low-hazard fuel-rich liquid monopropellant.

Consequently, in a first aspect the invention relates to anoxidizer-rich liquid monopropellant based on ADN or HAN.

A suitable engine has been disclosed in applicant's co-pendingapplications SE 1350612-6 and International patent application entitled“Dual mode chemical rocket engine and dual mode propulsion systemcomprising the rocket engine”.

In another aspect the invention relates to the use of the inventiveoxidizer-rich liquid monopropellant in a rocket engine in bipropellantoperation together with a fuel-rich liquid monopropellant based on ADNor HAN.

In yet an aspect the invention refers to a method of decomposing theinventive oxidizer-rich liquid monopropellant for the generation ofthrust, wherein the oxidizer-rich liquid monopropellant is injected intoa flow of hot fuel-rich gas obtained from the decomposition of afuel-rich liquid monopropellant, so that the oxidizer-rich liquidmonopropellant thereby is decomposed and combusted along with thefuel-rich gas, to increase the thrust.

In a related aspect the present invention relates to a method ofgenerating thrust, wherein the inventive oxidizer-rich liquidmonopropellant is injected into a flow of hot fuel-rich gas obtainedfrom decomposition of a fuel-rich liquid monopropellant, so that theoxidizer-rich liquid monopropellant thereby is decomposed and combustedalong with the fuel-rich gas.

The inventive oxidizer-rich liquid monopropellant can be used toimprove, in bipropellant operation mode, the performance of existingfuel-rich liquid ADN based monopropellants, such as LMP-103, LMP-103S,and FLP-106 (having the composition of 64.6% by weight of ADN, 23.9% byweight of water; and 11.5% by weight of MMF (N-methyl-formamide, akamono-methyl-formamide), and existing fuel-rich liquid HAN basedmonopropellants.

Using the inventive storable low-hazardous oxidizer-rich liquidmonopropellant together with a fuel-rich liquid monopropellant inbipropellant mode operation in a suitable dual mode chemical rocketengine, a propulsion system with comparable performance (i.e. in termsof total impulse for a given system mass) to the prior art dual modechemical propulsion systems can be achieved while avoiding the use ofthe prior art hazardous propellants.

A major advantage of the oxidizer-rich liquid monopropellant of theinvention is that it does not require a catalyst bed for thedecomposition of the oxidizer-rich liquid monopropellant. For thedecomposition of the fuel-rich monopropellant existing and well provencatalysts and catalyst beds currently used for the fuel-richmonopropellants can be used with the present invention.

The invention provides an enabling technology for substituting theconventional dual mode and bipropellant rocket propulsion systems usinghighly hazardous storable liquid propellants with a significantlyreduced hazard and environmentally benign alternative propellants systemwith comparable performance, and will also significantly reduce andfacilitate propellant handling and fuelling operations.

In the present invention the term “monopropellant” has been used todenote monopropellants which are composed of more than one chemicalcompound, such as LMP-103S, which thus could be regarded amonopropellant blend.

Further advantages and embodiments will be apparent from the followingdetailed description and appended claims.

BRIEF DESCRIPTION OF THE ATTACHED DRAWINGS

FIG. 1 shows a suitable dual mode chemical rocket engine wherein theinventive oxidizer may be used.

FIG. 2 shows a partial enlargement of means for injection 125 of theinventive oxidizer-rich monopropellant.

DETAILED DESCRIPTION

The inventive oxidizer-rich monopropellant comprises 70-90% of ADN orHAN, 0-10% ammonia, and balance water.

According to the present invention the inventive oxidizer-richmonopropellant is used for further combusting, in a second reactionstage, fuel-rich gasses obtained from combustion of a fuel-richmonopropellant, such as a conventional ADN-based or HAN-based liquidmonopropellant. The inventive liquid oxidizer-rich monopropellant isthus intended for use in bipropellant operation in a chemical rocketengine together with the fuel-rich liquid monopropellant.

As illustrated in FIG. 1 a suitable engine capable of operating inbipropellant mode may comprise a primary reaction chamber 130 for afuel-rich monopropellant, and a secondary reaction chamber 150 for thedecomposition of the inventive oxidizer-rich propellant, wherein theprimary reaction chamber is connected to the secondary reaction chamberso that fuel-rich gas from the decomposition of the fuel-rich oxidizerin the primary reaction chamber can flow into the second reactionchamber.

The inventive oxidizer-rich monopropellant is injected into thesecondary reaction chamber 150 via means for injection 125, e.g. aninjector.

In such engine the catalyst in the primary reaction chamber would be thelife limiting element of the thruster, when exposed to the reactivedecomposition and combustion species and operated at higher temperaturesthan their current design limits. By injecting the oxidiser-richmonopropellant into a secondary reaction chamber, wherein the fuel-richgas exiting the first reaction chamber is further combusted by means ofthe presence of the oxidizer-rich monopropellant, the temperature in thesecondary reaction chamber can be significantly increased, while thetemperature of the catalyst in the primary reactor can be keptessentially unaffected. Accordingly, existing and well proven catalystsand catalyst beds currently used for the specific fuel-richmonopropellant can be used in the primary reactor of such engine. Theprimary reactor can be based on similar reactor design as conventionalreactors for ADN based and HAN based liquid monopropellants,respectively, as currently used in corresponding liquid ADN and HANmonopropellant thrusters, respectively.

Thus, with the inventive oxidizer-rich monopropellant, existingtechnology can be used for combustion of the corresponding fuel-richmonopropellant, especially ADN monopropellant and HAN monopropellanttechnology, respectively.

It is generally preferred that the fuel-rich monopropellant blends, andoxidizer-rich monopropellant blends, respectively, be based on ADN.

In preferred embodiments the inventive oxidizer-rich monopropellant isto be used with a fuel-rich, liquid, aqueous ADN based monopropellant,such as e.g. LMP-103, LMP-103S, and FLP-106, especially LMP-103S.

The inventive oxidizer-rich monopropellant is formulated so as tomaximize, in bipropellant mode, the attainable combustion of fuel-richgasses exiting the first reactor. In principle, this means that theinventive oxidizer rich monopropellant will be formulated so that theoverall composition of the fuel-rich monopropellant and theoxidizer-rich monopropellant will correspond to the maximum obtainableI_(sp) of that overall composition.

According to calculations performed with NASA-Glenn Chemical EquilibriumProgram CEA2, operation of a chemical rocket engine with the inventiveoxidizer rich monopropellant in bipropellant mode would result in anadditional improvement of the specific impulse of up to 10% overLMP-103S, when used as a monopropellant only, which is about 10% lowerthan the specific impulse of the prior art bipropellant engines operatedon the highly hazardous conventional storable propellants, i.e. MMH andNTO. Furthermore, the density impulse of LMP-1035 and the inventiveoxidizer-rich ADN-blend combination will be up to 94% of the densityimpulse of the prior art bipropellant engine operated on conventionalstorable propellants.

Preferably, the oxidizer-rich monopropellant blend comprises 70-80% byweight of ADN or HAN Ammonia is preferably contained in an amount of1-10% by weight, more preferably 5-10% by weight, and especiallypreferred 5-8% by weight. The balance up 100% is water.

An especially preferred oxidize-rich ADN based monopropellant for use inthe dual mode chemical rocket engine comprises about 77% by weight ofADN, about 17% by weight of water and about 6% by weight of ammonia.

With reference to the engine 200 shown in FIG. 1 the operation and useof the inventive oxidizer will now be described in more detail by way ofexample.

The rocket engine 200 comprises one inlet port 101 for the fuel-richmonopropellant followed by a series redundant flow control valve 111 andpropellant feed tubes 121, and one inlet port 102 for the oxidizer-richpropellant followed by a series redundant flow control valve 112 andpropellant feed tube 122.

In bipropellant mode, the fuel-rich monopropellant LMP-103S, is injectedvia injector 110 into the primary reaction chamber 130, where thepropellant is thermo/catalytically decomposed (decomposition of ADNbased monopropellants have been disclosed in WO 02/095207) causing anexothermal reaction which produces heat up to about 1,600° C., and afuel-rich gas which flows into the secondary reaction chamber 150. Theinventive oxidizer-rich monopropellant (a composition of about 77% ADN,about 17% water and about 6% ammonia), is injected by means of a secondinjector 125 into the secondary reaction chamber 150 downstream of theprimary reaction chamber 130. A partial enlargement of injection means125 is shown in FIG. 2. In the secondary reaction chamber 150 theinventive oxidizer-rich monopropellant is atomized and decomposed thusgenerating a surplus of oxygen which mix in the secondary reactionchamber 150 with the fuel rich gases from the primary reaction chamber130. A secondary exothermal combustion takes place in secondary reactionchamber, wherein the stagnation gas temperature is significantly furtherincreased up to about 2,300° C. which enhances the performance of theengine in terms of fuel efficiency, i.e. specific impulse before theexhaust gases are accelerated through the nozzle 170 thus generatingthrust.

While described herein primarily with reference to a dual mode chemicalrocket engine, the inventive oxidizer-rich monopropellant could also beused in a similar chemical rocket engine designed for operation only inbipropellant mode.

1. An oxidizer-rich liquid monopropellant, consisting of: 70-90% byweight of an oxidizer selected from ammonium dinitramide (ADN) andhydroxyl ammonium nitrate (HAN); 0-8% by weight of ammonia; and balancewater.
 2. The oxidizer-rich liquid monopropellant of claim 1 comprising1-8% by weight of ammonia.
 3. The oxidizer-rich liquid monopropellant ofclaim 1, comprising 70-80% by weight of the oxidizer.
 4. Theoxidizer-rich liquid monopropellant of claim 1, wherein the oxidizer isADN.
 5. The oxidizer-rich liquid monopropellant of claim 4, comprisingabout 77% ADN, about 17% water and about 6% ammonia.
 6. Theoxidizer-rich liquid monopropellant of claim 1, wherein the oxidizer isHAN.
 7. A bipropellant combination comprising, stored separately, anoxidizer-rich liquid monopropellant according to claim 1, and, afuel-rich liquid monopropellant.
 8. The bipropellant combination ofclaim 7, wherein the fuel-rich liquid monopropellant is ADN-based orHAN-based.
 9. A rocket engine comprising the oxidizer-rich liquidmonopropellant of claim 1; and a fuel-rich liquid monopropellant basedon ADN or HAN.
 10. A method of decomposing an oxidizer-rich liquidmonopropellant for the generation of thrust, comprising injecting theoxidizer-rich liquid monopropellant of claim 1 into a flow of hotfuel-rich gas obtained from decomposition of a fuel-rich liquidmonopropellant, so that the oxidizer-rich liquid monopropellant therebyis decomposed and combusted along with the fuel-rich gas.
 11. A methodof generating thrust, comprising injecting an oxidizer-rich liquidmonopropellant according to claim 1 into a flow of hot fuel-rich gasobtained from decomposition of a fuel-rich liquid monopropellant, sothat the oxidizer-rich liquid monopropellant thereby is decomposed andcombusted along with the fuel-rich gas.
 12. The oxidizer-rich liquidmonopropellant of claim 2 comprising 5-8% by weight of ammonia.